Automatic airspeed engage/disengage

ABSTRACT

In an automatic flight control system (FIG. 1) an airspeed control engage function (84) is automatically engaged (136, FIG. 2; 244, FIG. 4) in response to airspeed above a threshold magnitude, such as 45 knots (88, FIG. 2) and will remain engaged (subject to a fault condition, 135) until the airspeed command (75) reaches a predetermined, insignificant magnitude (132). An airspeed error integrator 241 which accommodates the difference between a reference attitude and an attitude required for a reference airspeed, does not react to large airspeed errors as a consequence of pilot maneuvering due to pilot force on the control stick (35, 109) opening the input (252, FIG. 4) to the airspeed error integrator.

DESCRIPTION

1. Technical Field

This invention relates to automatic flight control systems, and moreparticularly to automatic engagement of an airspeed control system.

2. Background Art

In fixed wing aircraft, speed is normally controlled by engine thrust,and at any speed, aircraft pitch attitude (nose up or nose down) isutilized to control altitude. If speed is increased without adjustingpitch attitude, the aircraft will inevitably climb. In rotary wingaircraft (helicopters), the forward thrust required for increased speedis achieved by a combination of increased longitudinal cyclic pitch(nose down) with increased collective pitch (normal to the rotation ofthe main rotor blades).

Early automatic flight control systems for helicopters controlled speedsimply by means of attitude adjustments. If an automatic flight controlsystem includes both airspeed hold and pitch attitude hold functions, adesired airspeed (before engaging the airspeed hold function) isnormally achieved by pilot maneuvering of the longitudinal cyclic pitchcontrol, anticipating the necessary attitude for the desired speed (thespeed that will be achieved when the aircraft settles in a new attitudeat the conclusion of the pilot maneuver), and then automatic controlover the aircraft pitch attitude in response to an airspeed error signaldeveloped as the difference between the airspeed reference signalsynchronized to the airspeed at the conclusion of the maneuver, and thethen current airspeed.

In a typical airspeed hold system, the system is either engaged or notby the pilot. The system is operative at any airspeed (except below somethreshold level of reliable airspeed detection, such as 25 or 30 knots).If the pilot desires to maneuver the aircraft below or above the trimspeed, he can apply force to the pitch attitude controller (fore and aftmotion of the longitudinal cyclic pitch stick, in a helicopter), withoutcausing trim release, and thereby retain the airspeed reference so thatthe system will automatically regain the desired airspeed following themaneuver.

In a typical system, the longitudinal cyclic pitch full authority outerloop may either be fully disengaged, or have an integral path thereofdisengaged during the application of force to the control member.

In order to reduce pilot workload to thereby enhance aircraftperformance and safety in close maneuver operations (such as offshoredrilling and the like), it is desirable to provide a system which has anautomatically engaged airspeed control function. However, since airspeedis controlled by pitch attitude, accommodation must be made to transfercontrol from a pitch attitude synchronizer to an airspeed synchronizer,or to allow operation of both at the same time. Whenever both pitchattitude and speed are to be controlled, it is necessary to accommodatedeviations in pitch attitude from a reference pitch attitude, when thepitch attitude demanded for airspeed differs from the reference pitchattitude. The engagement and disengagement of airspeed should notprovide perturbations into the longitudinal cyclic pitch control system,and maneuvering of the aircraft by the pilot, while overriding thelongitudinal cyclic pitch control system, should not result in largeairspeed errors being retained at the conclusion of maneuvers, whichcould cause overshooting and oscillation about the desired airspeed, oncompletion of maneuvering.

DISCLOSURE OF INVENTION

Objects of the invention include provision of an automatically engagedand disengaged airspeed hold function.

According to the present invention, an airspeed control function in anautomatic flight control system is automatically engaged when theairspeed exceeds a predetermined threshold magnitude, and can becomedisengaged only when the airspeed command input to the automatic flightcontrol system falls below a relatively insignificant magnitude. Inaccordance further with the invention, the deviation between a referenceattitude and an attitude required to achieve the reference airspeed isaccommodated in an airspeed error integrator, which is inhibited fromtracking airspeed errors during pilot maneuver.

The invention permits smooth, unnoticed transition into and out of anairspeed hold function simply by synchronizing attitude at an airspeedabove a threshold airspeed (such as 45 knots), the airspeed controlfunction remaining engaged, even though the airspeed may thereaftertranscend below the threshold speed, until the airspeed command input tothe automatic flight control system is sufficiently low that thedisengagement will not provide a significant perturbation to the system.Utilization of an airspeed error integrator, to accommodate thedifference between any reference attitude resulting when attitude (andtherefore airspeed) is synchronized above the airspeed control thresholdspeed does not result in attitude or speed overshooting during pilotmaneuvers since the integrator is disengaged while force is applied tothe longitudinal axis of the cyclic pitch stick. The invention thereforeprovides essentially a phantom airspeed system which, although relatedto attitude, will hold speed (such as in the force of gusts and windshears), without causing perturbations during engagement ordisengagement, and without pilot induced errors. The system of thepresent invention may be implemented in automatic flight control systemsof a wide variety of types.

The invention may be practiced with analog, digital or computerizedsignal processing, utilizing only apparatus and techniques which arewell within the skill of the art, in the light of the specific teachingsrelating thereto which follow hereinafter. The foregoing and otherobjects, features and advantages of the present invention will becomemore apparent in the light of the following detailed description ofexemplary embodiments thereof, as illustrated in the accompanyingdrawings.

BRIEF DESCRIPTION OF DRAWINGS

FIG. 1 is a simplified schematic block diagram of an aircraft automaticflight control system in which the present invention may be implemented;

FIGS. 2 and 3 are simplified block diagrams of circuitry for providingcontrol signals for the automatic flight control system of FIG. 1; and

FIG. 4 is a simplified schematic block diagram of pitch attitudesynchronizing and beeping circuitry and of airspeed control circuitryfor the automatic flight control system of FIG. 1.

BEST MODE FOR CARRYING OUT THE INVENTION

Referring now to FIG. 1, a helicopter longitudinal cyclic pitch controlsystem for controlling the pitch axis attitude of a helicopter, withinwhich the present invention may suitably be implemented includes a pairof inner loop longitudinal cyclic pitch actuators 12, 13 which areconnected together by suitable linkage 14 and to a main rotor bladepitch angle swash plate mixer (not shown) by suitable linkage 15. Eachof the actuators is driven by a corresponding amplifier 16, 17 in anulling servo loop fashion. The amplifiers respond to error signals fromcorresponding summing junction 18, 19 which provide the amplifiers witha signal indicative of the difference between a pitch command signal ona related line 20, 21 and a signal provided on a related line 22, 23 bya corresponding actuator position sensor 24, 25 which is indicative ofthe achieved position of the actuator. When the actuators 12, 13 haveachieved positions corresponding to the signals on the lines 20, 21 theerror signal provided by the summing junctions 18, 19 to the amplifier16, 17 goes to zero, so the actuators will remain at rest until thesignals on the lines 20, 21 are changed (or drift showing up in thelines 22, 23).

The actuators 12, 13 are also connected by mechanical linkage 26 to acyclic pitch control stick 27 which is pivoted within a gimbal 28 forforward and aft motion against the operation of a trim position spring29. A pair of switches 31, 32 are disposed on the actuator 37 fordetecting motion of the stick 27 against the spring 29 in respectivedirections. Closure of either switch 31, 32 will provide a signal on acorresponding one of two lines 33 to cause an OR circuit 34 to provide apitch force signal on a line 35. In various embodiments, the OR functionprovided by the circuit 34 may simply be implemented by the relationshipof the switches 31, 32, as is known in the art.

The cyclic pitch stick 27 is connected by mechanical linkage 36 andspring 29 to a pitch outer loop trim actuator 37 which is driven throughpitch automatic shutdown circuits 38 and pitch pulser circuits 39 by asignal on a line 40 provided by a pitch outer loop integrator circuit41. These circuits are described in detail in a commonly owned copendingU.S. patent application entitled PULSED AIRCRAFT ACTUATOR, Ser. No.249,300, filed on even date herewith by Fischer et al. These circuitsserve to reposition the cyclic pitch stick 27 to a position indicativeof the actual commands being provided to the linkage 15 as a consequenceof motion of the actuators 12, 13. The pitch auto shutdown circuit 38provides a pitch outer loop shutdown signal on a line 42. The cyclicpitch stick 27 has a switch 44 that can be closed by a thumb or fingerso as to provide a trim release signal on a line 45. The stick 27 alsohas a "coolee hat" type of four-axis beeper switch 46 that can be movedforward or aft (or right or left) to provide beeper signals; in a systemof the type described, the beep signals are signals which force smallchanges in the attitude reference signals.

The pitch command signals on the lines 20, 21 are each provided by acorresponding summing junction 50, 51, which sum together correspondingpitch rate signals on lines 52, 53, pitch attitude and airspeed controlsignals on lines 54, 55 and outer loop compensation signals on lines 56,57. The outer loop compensation signals are provided by lag amplifiers58, 59 which are driven by the output of the pitch outer loop integratoron the line 40.

The signals on the lines 52-55 are applied to the pitch outer loopintegrator 41 in order to detect high pitch attitude demand changes. Thepitch attitude signals on the lines 54 and 55 are applied to a summingjunction 60, the output of which is applied to the pitch outer loopintegrator 41 on a line 61.

The signals on the lines 52, 53 are provided by differentiators 64, 65from gyro pitch signals on lines 66, 67 which are provided by the pitchaxis outputs of corresponding vertical gyros 68, 69. The signals on thelines 66, 67 are also compared with attitude reference signals in pitchattitude synchronizing and beep circuitry 70, 71. When the circuits 70,71 are synchronized, the reference follows (is made equal to) the signalon the corresponding lines 66, 67 indicative of actual pitch angle ofthe helicopter; when beeping is employed, the reference is forced toequal a greater or lesser pitch angle; when the circuits 70, 71 are notsynchronized, they provide pitch error signals on related lines 73, 74indicative of the variance between the actual pitch angle of thehelicopter and the desired helicopter pitch attitude. Logic circuits 72are connected with the pitch attitude synch and beep circuits 70, 71 tocontrol the operation thereof. In systems of the type disclosed herein,the signals on the lines 73, 74 are summed with a signal on a line 75 incorresponding summing junctions 76, 77, the resultant of which isapplied to a related limiter circuit 78, 79 so as to provide the pitchattitude and airspeed control signals on the lines 54, 55 limited to21/2 of total pilot authority. Thus, the short term automatic pitch axisinner loop control which can be provided by means of the actuators 12,13 is limited to ±5% (total, 10%) of total pilot authority.

To pitot-static airspeed system 80, of any suitable well known type,provides an airspeed signal on a line 83 to airspeed control circuits84. The airspeed control circuit 84 may also respond to the pitchattitude error signals on the lines 73, 74 so as to provide, over theline 75, increased gain in attitude control when the airspeed controlcircuitry 84 is engaged. The airspeed signal on the line 83 might beutilized in circuitry 86, which includes compare circuits and singleshot circuits (monostable multi-vibrators), or other signal transitiondetecting circuits, to provide signals on a plurality of lines 87-90indicative of airspeed being greater than 60 knots or 45 knots, andtransitions from above 45 knots to below 45 knots, respectively. Thiscircuitry may be of the type disclosed in U.S. patent application ofClelford et al, Ser. No. 176,832, filed on Aug. 8, 1980, or may be ofother types of hardware or software.

Referring now to FIG. 2, the logic circuits include a suitable voltagesource 100 which is connected through the beeper 46 so as to provide aforward beep request signal on a line 101 upon closure of beeper contact46f and to provide an aft beep request signal on a line 102 in responseto closure of contact 46a. An OR circuit 103 is responsive to a signalon either of the lines 101, 102 to provide a pitch beep signal on a line104. An inverter 105 is responsive to the signal on the line 104 toprevent the operation of an AND circuit 106 which otherwise isresponsive to the pitch force signal on the line 35 to provide a pitchstick signal on a line 107. The pitch stick signal on the line 107 isindicative of the fact that the pilot has moved the stick 27sufficiently in one direction or the other so as to close one of theswitches 31, 32, but has done so other than as a consequence of pushingthe beeper switch 46; this enables distinguishment between pitch forcesignals on the line 35 which really are as a consequence of overzealouspushing of the beeper switch 46, and pitch force signals on the line 35which are indicative of an intentional control input by the pilotthrough the pitch stick.

The pitch force signal on the line 35 is also applied to an OR circuit108 to provide a force delayed signal on a line 109. The force delayedsignal is indicative of the pitch force signal on the line 35, and isalso indicative of the fact that after the pitch force signal on theline 35 has been present for six seconds, it thereafter remains presentuntil six seconds after force is removed. This is achieved by having theOR circuit 108 connected to the set output of a bistable device 110which is set in response to an AND circuit 111 operable by the presenceof an output from a six second resettable single shot 112 concurrentlywith the pitch force signal on the line 35. The single shot 112 is inturn initiated (at its set input) by the pitch force signal on the line35 and will, when initiated, initially lose the signal on a line 113 atits complementary output, but after six seconds will regain the signalon the line 113 and apply it to the AND circuit 111, unless the singleshot 112 is reset prior thereto by a signal from an inverter 114 as aconsequence of the pitch force signal no longer being present on theline 35. Thus, if the pitch force signal on the line 35 lasts for lessthan six seconds, the bistable device 110 will not become set. But if itlasts for more than six seconds, the bistable device 110 will becomeset. When the pitch force signal disappears from the line 35, theinverter 114 starts another six second single shot 115 and causes thesignal on its complementary output to disappear so there is no signal ona line 116 to the reset input of the bistable device 110. But after theexpiration of the six second pulse, the complementary output will againachieve a signal on the line 116, the rise of which will cause resettingof the bistable device 110. Thus the OR circuit 108 will provide a forcedelayed signal during the presence of pitch force, and if pitch forcelasts at least six seconds, it provides the force delayed signal for sixseconds after termination of the pitch force signal on the line 35. Asis described more fully hereinafter with respect to FIG. 4, thisprevents the airspeed control circuits 84 from integrating airspeederror during pilot input and for six seconds after he completes amaneuver.

The force delayed signal on the line 109 is also utilized to signify asignificant pilot input in the pitch axis which may (if in the aftdirection) cause the airspeed of the helicopter to drop below that atwhich an automatic airspeed hold function (described more fullyhereinafter) is to be engaged. For instance, if the pilot induces asignificant nose-up or slow-down maneuver, the airspeed may fall below45 knots (used in the example herein to be exemplary of the airspeedhold function). But, custom or government regulation frequently requiresthat removal of stick force (following a maneuver commenced while theairspeed hold function is engaged) should cause the original, referenceairspeed to be recovered automatically. Therefore, the airspeed holdfunction should not become disengaged if it was engaged prior todropping below the critical airspeed. The force delayed signal on theline 109 is therefore fed to an AND circuit 120 which is operativewhenever there is a signal on a line 88 indicating that airspeed isgreater than 45 knots. The AND circuit causes a bistable device 122 tobecome set thereby providing a signal to an OR circuit 123 that is alsoresponsive to the airspeed greater than 45 knots signal on the line 88.This provides an airspeed engage enable signal on a line 124 wheneverthe airspeed is above 45 knots, or, having been above 45 knots whenforce is applied to the stick. The bistable 122 will remain set untilthe airspeed recovers to substantially the initial airspeed as indicatedby a signal input to an AND circuit 125 which is generated by a comparecircuit 126 which includes a -3 knots reference. The circuit 126 isresponsive to a signal on a line 127 to indicate when the airspeed error(the difference between the reference airspeed and actual airspeed) iswithin -3 knots. Thus the signal on line 126a indicates that theaircraft has recovered a speed which is not more than 3 knots lower thatthe original reference airspeed. As a consequence of the foregoing, thebistable 122 will become set when airspeed is greater than 45 knots ifforce is applied by the pilot, and will thereafter remain set untilforce is removed and airspeed is within 3 knots of reference airspeed.Thus the airspeed engage enable signal is present on the line 124whenever the airspeed is greater than 45 knots, or when the airspeed hasbeen greater than 45 knots, a stick force has been applied, and theaircraft has not yet regained all but 3 knots of the original airspeed.

The airspeed engage enable signal on line 124 is applied to an ORcircuit 130 which is also responsive to a signal on a line 131 from acompare circuit 132 that is responsive to the airspeed command signal onthe line 75. The compare circuit 132 has a reference voltage in it whichis equivalent to ±1% of full pilot authority; therefore the signal onthe line 131 will be present unless the airspeed command on the line 75is essentially nil. This avoids having the airspeed controls disengage(by transition in speed to less than 45 knots) while there is a largeairspeed command, which could cause a jump in the pitch command. The ORcircuit 130 feeds an AND circuit 134 that is blocked by an inverter 135whenever the pitch outer loop shutdown signal is present on the line 42.Therefore, whenever the pitch outer loop is operative, and the airspeedis greater than 45 knots, there will be an airspeed engaged signal on aline 136. The airspeed engage signal on line 136 will continue to bepresent even though the airspeed drops below 45 knots if force isapplied to the stick, and this condition of forcing the airspeed engagedto remain present below 45 knots will continue until the airspeed erroris less than minus 3 knots (the present airspeed is within 3 knots ofthe original airspeed before the force was applied).

At the bottom of FIG. 2, the trim release signal on the line 45 isapplied to a 7/10 second single shot 137, the output of which enables anAND circuit 138 to provide an initial trim release signal on a line 139during the first 7/10 second of the appearance of the trim releasesignal on the line 45. When the 7/10 of a second has timed out, theoutput of the single shot 137 will disappear, blocking the AND circuit138 to end the initial trim release signal on the line 139.

Referring now to FIG. 3, an OR circuit 143 responds to the airspeed uptransition signal on the line 89, the pitch beep signal on the line 104,or the trim release signal on the line 45 to provide a signal on a line144 that will be passed by an AND circuit 145, provided an inverter 146is not activated by a signal on a line 147. The AND circuit 145 providesa signal on a line 148 that causes an OR circuit 149 to generate a synchrequest signal on a line 150. The signal on the line 148 is applied toan inverter 155 so that when the signal disappears, the inverter 155will energize the set input of a 25 second single shot 156, the trueoutput of which on a line 157 is applied to the OR circuit 149. Thismeans that, in the usual case, once the synch request signal isgenerated on the line 150 by a signal on the line 148, that signal willbe maintained for 25 seconds after the signal on the line 148disappears, due to the signal on the line 157 which is present for 25seconds after the signal disappears from the line 148. The purposes forthis are described more fully hereinafter.

The signal on the line 147 is provided by an OR circuit 160 in responseto the pitch stick signal on the line 107 (indicating that the pilotintends to apply force to the stick), or the pitch outer loop shutdownsignal on the line 42 (indicating that there will no longer be any pitchouter loop inputs to the mixer), or in response to a signal on a line161 from an inverter 162 indicative of the absence of an airspeed engageenable signal on the line 124. Thus the signal on the line 147 isindicative of the airspeed being or soon to become disengaged, the pilotinducing an intentional input, or the loss of the pitch outer loopinput. The signal on the line 147, being applied to the inverter 146 andto the reset side of the single shot 156 will prevent the OR circuit 149from presenting the synch request signal on the line 150. If a synchrequest is in process, it will terminate in response to any of thesesignals. The main function of the synch request signal is to provide,through an OR circuit 164, an airspeed synchronizing signal on a line165. This signal can also be presented in response to signals on thelines 42 or 161, or in response to the trim release signal on the line45. As is described hereinafter, the airspeed synch signal on the line165 causes the airspeed reference signal to be equal to current airspeedat all times when it is present. Thus whenever the pilot presses thetrim release, all of the cyclic pitch autopilot functions (airspeed,pitch attitude and roll attitude) are terminated and new references areestablished as long as the trim release is pressed. Whenever automaticairspeed hold is (or is about to be) ended, the airspeed error is forcedto zero (meaning no airspeed input to the system) by the airspeed synchsignal on the line 165. There is no airspeed hold permitted when thepitch outer loop is shut down. Otherwise, the OR circuit 143 generallycauses a 25 second airspeed synch signal unless it is previouslyterminated by the OR circuit 160, for any operation of pitch beep ortrim release (no matter how short) or upon a transition from below 45knots to above 45 knots, and is held for 25 seconds after thetermination of such events.

An airspeed integrator reset signal is generated on a line 167 by an ORcircuit 168 in response to any of the three signals on lines 42, 161, or45, or in response to an auto-synch signal on a line 169. The auto-synchsignal on the line 169 is a pulse provided by a half-second single shot170 whenever there is an output from a window comparator 171 thatcompares the output of the airspeed integrator on a line 172 with plusand minus reference voltages equivalent to 8% of full pilot authority.As described hereinafter with respect to FIG. 4, whenever the airspeedintegral gain path is providing a signal equal to ±8% of pilotauthority, it causes a slewed reduction in the integrator output for ahalf-second, along with a commensurate reduction in the attitudesynchronizer integrator. This will reduce the opposing authorities ofthe airspeed and attitude inputs to the system while retaining thebalance between them.

The auto-synchronizing signal on the line 169 is also applied to an ORcircuit 175 to cause the pitch attitude synchronizer to have theequivalent half-second slewed reduction in the pitch attitude referencevalue. The OR circuit 175 will activate a half-second, resettable singleshot 176 which therefore provides a half-second pulse on a line 177,unless that pulse is terminated by application of the pitch stick signalon the line 107 to the reset input of the single shot 176. Thus, in thenormal course of events, the generation of the half-second auto-synchpulse on the line 169 will commensurately cause a half-second pitchattitude synch pulse on a line 178 from OR circuit 179 that isresponsive to the single shot 176. In addition, the pitch attitude synchsignal on the line 178 is continuously present during trim release asindicated by the signal on the line 45. And, half-second pitch attitudesynch pulses can also be provided anytime there is a transition fromabove 45 knots to below 45 knots as a consequence of the airspeed downtransition signal on the line 90 being applied to the OR circuit 175.Also, whenever the synch request signal on the line 150 terminates, aninverter 181 will operate the OR circuit 175 to provide a pitch attitudesynch signal on a line 178. Thus a trim release signal cansimultaneously provide the pitch attitude synch signal on a line 178,cause a commensurate signal on the line 148 to generate a synch requestsignal on the line 150, cause single shot 156 to extend the synchrequest signal on the line 150 for 25 seconds after disappearance of thetrim release signal, so that the airspeed remains synchronized for 25seconds after the pitch attitude has ceased to be synchronized, and whenthe 25 seconds are up, the absence of the synch request will cause onelast pulse of pitch attitude synch signal on the line 178. Theutilization and purposes of these signals are described more fully withrespect to FIG. 4 hereinafter.

Referring to the bottom of FIG. 3, a plurality of comparators 183-186are responsive to the pitch attitude and airspeed signals on the lines54, 55 to provide corresponding signals on lines 190-193 whenever thepitch attitude and airspeed command signals on the lines 54, 55 aregreater than ±21/2% of pilot authority. Thus, if both of the pitchattitude and airspeed signals on the lines 54, 55 are in excess of+21/2% of pilot authority signals will be present on the lines 190 and192 which will cause an AND circuit 195 to operate an inverter 196 andtherefore block an AND circuit 197. On the other hand, if both of thepitch attitude signals on the lines 54, 55 have magnitudes equivalent tomore than -21/2% of pilot authority, there will be signals present onlines 191 and 193 which will cause an AND circuit 200 to operate aninverter 201 so as to block an AND circuit 202. This is a beep inhibitfunction that prevents any attempt to beep in the same direction as anexisting saturating pitch attitude and airspeed command, so that thereference will not continually lead the system beyond the capability ofthe system to respond in view of the 21/2% limiters 78, 79 (FIG. 1). Asan example, if a helicopter takes off and then noses down to gain speed,spurious aerodynamic effects could cause the helicopter to retain arelatively more level attitude than that which is being commanded to themaximum 5% pilot authority. Should the pilot attempt to beep the nosefurther over the greater airspeed, the inner loop pitch actuators willnot be able to command additional nose down attitude over a relativelylong term of several seconds or more. Any attempt to beep the nose overat a greater rate would simply build up the attitude reference voltagebeyond what the aircraft can respond to. As the aircraft tends torecover, it can nose over (tuck under) in an undesirable overshootingfashion. Thus, whenever the pitch channels of the aircraft are drivingthe inner loop to the maximum electrical limit, no beeping (no drivingof the reference) is permitted in that direction, to preclude anyovershoot command condition to develop.

Referring now to FIG. 4, the pitch attitude synch and beep circuits 70comprise a unique modification of the typical integral feedback circuitof the type known in the art, such as that disclosed in theaforementioned Clelford patent. Specifically, a summing junction 206subtracts a pitch attitude reference signal on a line 207 from the pitchaxis output of the first vertical gyro on the line 66 so as to providethe pitch error signal on the line 73. The reference signal on the line207 is established and held by an integrator 208, the input of which iscontrolled by a contact 209 of a normally open relay which closes whenits coil 210 is energized. When the contact 209 is closed, and provideda switch 210a is energized, then the error signal on line 73 is fed backthrough a variable gain amplifier 211 to the input of the integrator208. Thus, depending upon the gain of the amplifier 211 and the lengthof time during which it is connected to the input of the integrator 208,the integrator 208 will integrate until it has an output voltage on theline 207 which is equal to the pitch voltage on the line 66, so that thepitch error signal on the line 73 is zero, and no further integrationoccurs. This is called synchronizing. Synchronizing occurs by virtue ofthe pitch attitude synch signal on the line 178 causing an OR circuit212 to operate the coil 210 and close the contact 209, as well asoperating the switch 210a so as to interconnect the amplifier 211 withthe integrator 208.

The amplifier 211 consists of an operational amplifier 214, the gain ofwhich is the balance between an input resistor 215 and feedbackresistance. Normally, the only feedback resistance is provided by aresistor 216. But in certain circumstances, other resistors are placedin parallel therewith so as to lower the feedback resistance and therebysignificantly decrease the gain. For instance, a resistor 217 is placedin circuit by closing a switch 218 in response to the initial trimrelease signal on the line 139. Thus, when trim release is the signalthat (through the OR circuit 179, FIG. 3) causes the pitch attitudesynch signal on the line 178, the amplifier 211 will have a relativelylow gain during an initial 7/10 of a second due to the appearance of theinitial trim release signal on the line 139 causing the switch 218 toinsert the resistance 217 to significantly lower the feedbackresistance. After 7/10 second, the signal on line 139 ends and the highgain is restored. Since the effective time constant of the integrator208 is an inverse function of the gain of the amplifier 211, the timeconstant during trim release will initially be relatively large comparedto what it is after 7/10 of a second. For instance, the gain of theamplifier 211 may be adjusted so that an initial time constant is 500 MSand after 7/10 of a second the time constant may decrease to 16 MS. Thepurpose of this function is so that, when trim release is depressed, thereference voltage and therefore the error voltage will initially changerelatively slowly for a smooth inner loop command transition. But afterthe initial time frame (0.7 seconds), the synchronizing circuit respondsvery quickly to changes which occur as a consequence of variations inthe pitch attitude of the aircraft showing up as changes in the pitchaxis voltage on the line 66. When the trim release signal ends, thesmall time constant synchronizing will be accurately reflecting thecurrent pitch angle of the aircraft, causing a nearly nil pitch errorsignal on the line 73 when the trim release signal is removed. In amanner described more fully hereinafter, a different resistor 220 may beplaced in parallel with the resistor 216 by operation of a switch 221 inresponse to the auto-synch signal on the line 169. Instead of alteringthe gain of an input amplifier 211, the same effect could be had byselectively switching in different feedback capacitance across anintegrating amplifier (within the integrator 208), as is well known tothose skilled in the art.

Another function of the synchronizer is to allow slow changes in thereference signal on the line 207 as a consequence of beeping (providinggradual adjustments to) the input to the integrator 208. This may beachieved by the pitch beep signal on the line 104 causing the OR circuit212 to energize the coil 210 and close the contact 209, withoutoperating the switch 210a. Then, small positive or negative DC voltagesfrom respective sources 224, 225 may be applied through the contact 209to the input of the integrator 208 by closure of a forward deep commandswitch 226 or an aft beep command switch 227 in respective response tothe forward beep command signal on the line 203 or to the aft beepcommand signal on the line 204. As described with respect to thecircuitry 183-204 in FIG. 3, the integrator 208 is not allowed to bedriven further in a given direction when the pitch error signal on theline 73 is such as to be driving the 21/2% limiters 78, 79 (FIG. 1) tosaturation. Therefore, even though a beep switch may be closed causingthe forward or aft beep request signal on the line 101, 102, (FIG. 3),the beep command signals will not be applied to either of the switches226, 227 so that further error is not incurred. This avoids building upthe reference signal way beyond that which is desired as a consequenceof the inability of the aircraft to assume the desired attitude in areasonable time due to aerodynamic effects.

The airspeed circuitry 84 includes airspeed synchronizing circuitrywhich is a simple version of that described with respect to the pitchattitude synch and beep circuit 70, hereinabove. A summing junction 230provides the airspeed error signal on the line 127 as the differencebetween an airspeed reference signal on a line 231 and the airspeedsignal on the line 83. The airspeed reference signal on the line 231 isprovided by an integrating amplifier 232 which is the equivalent of thecombination of the amplifier 211 and the integrator 208 in the upperportion of FIG. 4, or which may simply be an integrator havingcapacitive feedback along with a resistive input, as is well known inthe art. The integrating amplifier 232 is connected to the airspeederror signal on the line 127 whenever a switch 233 is operated by theairspeed synch signal on the line 165. When the switch 233 is closed,the integrating amplifier 232 will, in dependence upon the gain and timeconstant thereof, provide a reference signal on a line 231 that willcause an airspeed error signal on the line 127 which is essentially nil,thereby to synchronize the airspeed reference to essentially the currentairspeed of the aircraft. For no reference leakage, a relay may be usedin place of the switch 233.

The airspeed error is fed to a proportional gain path 240 and anintegral gain path (241) to a summing junction 242, the output of whichon a line 243 is connected through a switch 244 to comprise the airspeedcommand signal on the line 75, whenever airspeed is engaged as indicatedby the signal on a line 136. The proportional gain path 240 consists ofan amplifier 246 which feeds a ±3.7 knot limiter 247. This allows arelative high gain near the trim airspeed, without overshooting inresponse to large airspeed errors. The integral path 241 includes anintegrator consisting of an amplifier 248 with a resistive input 249 anda feedback capacitor 250. When a switch 252 is closed, it causes theairspeed error signal on the line 127 to be fed through a ±2 knotlimiter 253 to the integrator; the limiter prevents high airspeed errorsignals from building up too rapidly in the integrator and permittingthe contribution of the integral path 241 to become so great as toresult in the aircraft overshooting the desired airspeed. Thus, if theairspeed signal on the line 83 changes dramatically as a consequence ofa heavy gust or as a consequence of an input by the pilot, without thelimiter 253, the error built up in the integrator could causeovercompensation for speed, resulting in a slow oscillation in airspeed(and pitching of the aircraft) as the airspeed settles down thereafter.For the same reason, whenever there are pilot inputs the force delayedsignal on the line 109 causes an inverter 255 to open the switch 252(which could be a relay) so that there will be no input to theintegrator, and the errors induced by the pilot input will notcontinuously build up in the integrator over a long period of time. Theairspeed error on the line 127 which builds up during pilot maneuverswill drive the aircraft back to the desired airspeed at the end of pilotmaneuvers; so long as the buildup within the integral path 241 isinhibited, only minimal overshoot or consequent oscillation in airspeedwill occur as the reference airspeed is regained following pilotmaneuver. If the pilot maneuver is long (in excess of 6 seconds), theintegrator is held off for 6 seconds after end of pilot input to allowthe airspeed to reduce significantly, so as not to drive the integratorwith the initial high error, and to reduce the time over which the erroris integrated.

As described with respect to FIG. 1, the airspeed command signal on theline 75 (when the airspeed is engaged) is summed with the attitude erroron the line 73 (as well as the attitude error on the line 74).Therefore, there is an interrelationship between the pitch attitudesynch and beep circuits 70 and the airspeed control circuits 84. In FIG.4, it is seen that the only way the pilot can beep to a desired airspeedis by beeping attitude.

If a heavy, long trim head-on gust (for instance) reduces airspeed, theairspeed error will build up in the integrator 241, causing an attitudechange to regain the airspeed. The attitude change results in anattitude error on line 73. These opposing effects can build up to such apoint that their inputs are equal and oppositely saturated. In such acase, retrimming of the whole system could be required should the pilotsense that he lost airspeed retention capability. To avoid that, theauto-synch signal on the line 169 is provided as described hereinbeforewhenever the airspeed integrator output reaches ±8% of full pilotauthority. In FIG. 3, the auto-synch signal on the line 169 causes theairspeed integrator reset signal on the line 167 which is applied (inFIG. 4) to a switch 258. The switch 258 causes a resistor 259 to beconnected in parallel with the capacitor 250 and cause the capacitor tobleed down with an equivalent time constant of one-half second.Simultaneously, the auto-synch signal on the line 169 applied to theswitch 221 causes the gain of the amplifier 211 to be reduced byparalleling of the resistor 220 with resistor 216, so that thecombination of the amplifier 211 and the integrator 208 will partiallysynchronize the pitch attitude error on line 73 with a one-half secondtime constant as well. Thus the pitch attitude error signal on the line73 is reduced commensurately with the output of the airspeed integratorpath 241 at an equal and opposite rate. This transpires only for ahalf-second because the auto-synch signal on the line 169 is generatedby the half-second signal shot 170 (FIG. 3) and the pitch attitude synchsignal is generated by the half-second single shot 176 in response tothe auto-synch signal on the line 169. Since these circuits are operatedwith a half-second time constant for one-half second, the referencevoltages will be reduced by 63% of their original value whenever thathappens. The equal percentage reduction balances out even though thesynch circuit 70 provides only half of the opposing gain, the pitchattitude synch and beep circuit 71 (FIG. 1) providing the other half.Thus, the airspeed integrator 241 will reduce from ±8% of full pilotauthority to about 3% of full pilot authority as each of the pitchattitude errors (208) reduce from ±4% of pilot authority to about 1.5%of pilot authority. Therefore saturated, equal but opposite operation isautomatically avoided, and full control can readily be retained eventhough the introduction of airspeed errors to trim pitch attitude maycause large pitch attitude errors subsequent to locking onto airspeed.

In order to provide increase gain for additional dynamic stability inthe pitch channel when the airspeed control is engaged, a pair ofamplifiers 262, 263 apply inputs to the summing junction 242 from thecorresponding pitch error signals on the lines 73 and 74.

One feature of the apparatus described is that airspeed control isautomatically engaged and automatically disengaged as a function ofairspeed. In order to ensure that (other than cases when the outer loopshuts down) the airspeed control does not disengage disruptively, theairspeed control is not allowed to become disengaged except when itsoutput to the pitch channels is very small (as indicated by the 1%comparator 132 in FIG. 2). Thus if the pilot purposely lowers the speedof the aircraft below 45 knots and attempts to trim up on a newairspeed, resynchronization of the pitch attitude channel will causeimmediate resynchronizing of the airspeed and pitch attitude errors andresetting of the airspeed integrator (due to the circuitry at the top ofFIG. 3). When this synchronizing and resetting occurs, the contributionto the inner loop by the airspeed command will be essentially nil sincethe pitch attitude error on line 73 (FIG. 4) will supply a smallcontribution through the amplifier 262 and the airspeed error beingessentially nil will supply a small contribution through theproportional path 240 and the airspeed integrator 241. Then, theairspeed engage signal on the line 136 can disappear, opening the switch244 and returning the aircraft to simple attitude control, rather than acombination of airspeed and attitude control.

The invention is implemented simply, as is illustrated in FIGS. 2 and 4.In FIG. 2, airspeed being greater than 45 knots will automatically causethe airspeed engage enable signal on the line 124 which automaticallycauses the airspeed engaged signal on the line 136. Once engaged, unlessthe pitch outer loop is totally shut down, the airspeed will remainengaged until the airspeed falls below 45 knots and the airspeed commandis less than 1% of the pilot authority as indicated by the comparator132. Beeping to an airspeed below 45 knots, or trim release below 45knots, will synch the airspeed, and drive the airspeed command to zero,thereby to disengage. In addition, whenever airspeed is engaged,additional gain in the pitch attitude channel is achieved by means ofthe amplifiers 262, 263 (FIG. 4). And, the discrepancy between thereference attitude established by the pilot synchronizing both attitudeand airspeed in anticipation of a desired airspeed, and the attitudenecessary to give the desired airspeed so synchronized, is accommodatedin the airspeed integrator 241. However, whenever the pilot maneuvers,the sensing of force on the stick will cause the input (switch 252) ofthe airspeed integrator to be open, so that large airspeed errorsresulting from maneuvering will not build up in the integrator. And, ifthe pilot maneuvers by force for more than six seconds, the forcedelayed signal on the line 109 will continue to block the input to theintegrator 241 for six additional seconds after force is removed, toallow attitude to settle down. Therefore, buildup of airspeed errorswhich could cause overshoot and oscillation, following the release offorce on the cyclic pitch stick, are avoided by blocking the integrator.Thus, an essentially phantom airspeed system, which comes on lineautomatically above 45 knots as a result of pilot synchronizing or thetransition at a given attitude and airspeed, provides no perturbationswhen automatically engaged, and substantially no perturbations whendisengaged below the threshold airspeed with a minimum airspeed commandbeing provided.

The foregoing description is in simplified block form, the detailedcircuitry being described with respect to simple positive logicutilizing either relay contacts or switches to open and close certainpaths, summing junctions (which are understood to be combinations ofresistors at the correct inverting and noninverting inputs of suitableamplifiers) single shots which may require reset dominance or may notneed one, bistable devices, and the like. Many of the foregoingfunctions can obviously be achieved in a simpler fashion by using moretrue and complement outputs and fewer inverters; in many instances thepositive logic disclosed may readily be reworked into inverting logic tobe more suitably applicable to available hardware chips. The descriptionis, therefore, principally in terms of function achieving blocks, and itshould be understood that numerous variations may be utilized forachieving the same or equivalent functions and combinations of functionswithin the skill of the art. In addition, the functions of the foregoingapparatus (other than the mechanical functions and those functions whichdirectly interface with the mechanical functions) may readily beimplemented by utilization of a suitably programmed digital computer.The conversion of the discrete and analog functions described herein todigital functions performed by suitable software in a computer is wellwithin the skill of the art, particularly in the light of the teachingsof equivalency set forth in a commonly owned copending U.S. patentapplication, Ser. No. 176,832, filed on Aug. 8, 1980 by Clelford et al.

The invention may be practiced in automatic flight control systemshaving single channels of inner loop or of outer loop, dual channels ofinner loop or of outer loop, or more channels of either, in variouscombinations. The exemplary conditions, magnitudes, durations andrelationships may of course be varied to suit any usage of theinvention. Aspects of the invention may be practiced in automaticcontrol of various functions, in addition to the illustrative functionsdescribed herein.

Similarly, although the invention has been shown and described withrespect to an exemplary embodiment thereof, it should be understood bythose skilled in the art that the foregoing and various other changes,omissions and additions may be made therein and thereto, withoutdeparting from the spirit and the scope of the invention.

We claim:
 1. A system for positioning pitch-attitude-controllingaerodynamic surfaces of an aircraft, comprising:control means responsiveto force applied by a pilot to position said aerodynamic surfaces; pitchattitude means for providing an attitude signal indicative of the actualattitude of the aircraft in its pitch axis; airspeed means for providingan airspeed signal indicative of actual aircraft airspeed; actuatormeans responsive to a command input signal applied thereto forpositioning said aerodynamic surfaces; and signal processing means,responsive to said pitch attitude means and said airspeed means forproviding a pitch reference signal indicative of the pitch attitudedesired for the aircraft, for providing an airspeed reference signalindicative of a desired airspeed for the aircraft, for selectivelyproviding an airspeed engage signal indicative of the fact that controlof said actuator means includes commands for attaining said desiredairspeed, for providing an attitude error signal indicative of thedifference between said pitch reference signal and said actual attitudesignal, for providing an airspeed error signal indicative of thedifference between said airspeed signal and said airspeed referencesignal, for alternatively providing to said actuator means a pitchattitude command signal indicative of a desired change in aircraft pitchattitude in response to said attitude error signal when said airspeedengage signal is not present, or in response to both said attitude errorsignal and said airspeed error signal when said airspeed engage signalis present; characterized by said signal processing means comprisingmeans for providing said airspeed engage signal in response to saidairspeed signal indicating an airspeed in excess of a predeterminedthreshold magnitude, and for thereafter continuing to provide saidairspeed engage signal even if said airspeed signal indicates anairspeed less than said predetermined magnitude in concurrent responseto the presence of said airspeed engage signal and said pitch attitudecommand signal indicating a desired change in aircraft pitch attitude inexcess of a preestablished magnitude.
 2. A system according to claim 1including means operable by a pilot for providing an attitude referenceadjusting signal;characterized by said signal processing meanscomprising means responsive to said attitude reference adjusting signalto cause said airspeed reference signal to be substantially equal tosaid airspeed signal, whereby said pitch command signal will become lessthan said preestablished magnitude when the attitude of said aircraft issuch as to provide said actual attitude signal substantially equal tosaid pitch reference signal.
 3. A system according to claim 1characterized by: means responsive to said control means for providing aforce signal indicative of the pilot applying force to said controlmeans; andsaid signal processing means comprising means for providing anairspeed command signal as a proportional and integral function of saidairspeed error signal, for providing said pitch attitude command signalin response to said attitude error signal and said airspeed commandsignal when said airspeed engage signal is present, and for holding saidintegral function constant in response to the presence of said forcesignal, whereby airspeed error signals created by pilot maneuvers arenot included in said integral function.
 4. A system according to claim 1characterized by said signal processing means comprising means forproviding said airspeed reference signal substantially identical to saidairspeed signal in response either to said airspeed signal initiallyindicating an airspeed in excess of said predetermined magnitude or tothe absence of said airspeed engage signal.